1. Field of the Invention
The present invention relates to ventilating a high pressure turbine in a two-spool turbomachine such as an airplane turbojet, and it relates more particularly to ventilating a high pressure turbine disk.
2. Description of the Related Art
Two-spool turbomachines have a high pressure turbine arranged at the outlet from a combustion chamber in order to extract energy from a stream of gas ejected by the combustion chamber and impart rotary drive to a high pressure compressor located upstream from the combustion chamber in order to feed said chamber with air under pressure. Such turbomachines also include a low pressure turbine arranged downstream from the high pressure turbine in order to extract additional energy from the gas stream and impart rotary drive to a low pressure compressor that is arranged upstream from the high pressure compressor.
The high pressure turbine generally comprises a disk located at the outlet from the combustion chamber and carrying blades that are driven in rotation by the stream of gas ejected by said combustion chamber, the disk being surrounded by a stator element such as a sectorized ring in order to seal the flow section for gas through the turbine.
Because of the high temperatures reached by the combustion gas, the stator sealing ring and the rotor disk are subjected to high levels of thermal stress of a kind that causes these components to expand.
The disk presents relatively high mass and therefore responds more slowly than the sealing ring to variations in the temperature of the gas, which variations of temperature are caused by variations in the operating speed of the turbomachine, thereby giving rise to differential thermal expansion, particularly since the disk is less exposed to the combustion gas than are the blades that it carries and the sealing ring of the stator.
Such differential thermal expansion gives rise to variations in the clearance at the tips of the blades during various operating stages of the turbomachine, thus making it necessary to provide relatively large amounts of clearance, to the detriment of turbine performance.
Furthermore, temperature within the disk is not uniform, in particular between its radially outer periphery carrying the blades, which are in contact with the combustion gas, and its hub, which is spaced apart from the combustion gas.
Temperature gradients in the disk shorten its lifetime and make it necessary to use a disk that is relatively thick and massive, which goes against the attempts at achieving weight savings that are inherent to designing such turbomachines.
In order to limit those drawbacks, the disk is generally ventilated by air bled from upstream in order to heat it when speeds are increasing so as to accelerate its thermal expansion, and in order to cool it when speeds are decreasing so as to accelerate its contraction.
The blades of the disk generally benefit from a dedicated ventilation circuit that bleeds air from the combustion chamber end in order to convey said air via injectors into an annular cavity formed immediately upstream from the disk and communicating with ventilation circuits formed inside the blades.
The hub of the disk receives ventilation air that is generally bled from a stage of the high pressure compressor, and that flows downstream, e.g. along a cylindrical shroud or sheath that extends axially from the above-mentioned stage of the compressor and that defines an annular cavity that is radially inside the above-mentioned cavity extending downstream as far as the disk of the high pressure turbine.
Nevertheless, air taken from the high pressure compressor for ventilating the hub of the disk does not have the same temperature as the air that is bled from the combustion chamber end in order to ventilate the blades of said disk, and it follows a path that is considerably longer. During a change of speed, the air for ventilating the hub of the disk thus experiences temperature changes with a delay relative to the air for ventilating the blades and relative to the combustion gas.
This makes it difficult to control the clearance at the blade tips and makes it necessary to provide relatively large amounts of clearance of a kind that penalizes performance of the turbine, in order to limit the risk of premature wear of the blades and of the sealing ring surrounding them.
Furthermore, this is harmful to achieving a satisfactory reduction of temperature gradients in the disk of the turbine.
In addition, the air bled from the combustion chamber end for ventilating the blades of the high pressure turbine is at a pressure that is higher than the pressure of the air bled from the high pressure compressor for ventilating the hub of the disk of said turbine.
The blade ventilation air passes via the annular cavity connected to the internal ventilation circuits of the blades, and it therefore applies pressure to a radially outer portion of the upstream flank of the disk, while the hub ventilation air flows through the radially inner cavity on both sides of the hub of the disk.
This gives rise to the disk being subjected to unbalanced pressures, thereby inducing downstream axial thrust on the disk and making the turbomachine more difficult to control.